Stiffening member for epicyclical gear system housing assembly

ABSTRACT

A planet gear housing assembly is disclosed in an epicyclical gear system of a gas turbine engine having an engine casing. The planet gear housing assembly comprises an aft planet carrier assembly, a forward planet carrier assembly, and a plurality of planet gears. An aft flange of the aft planet carrier assembly is coupled to the engine casing to define a first torsional stiffness. A forward flange of the forward planet carrier assembly is coupled to the aft flange to define a second torsional stiffness. The second torsional stiffness may be between 60% and 80% of the first torsional stiffness. The planet gear housing assembly may further comprise a stiffening member positioned between the forward planet carrier assembly and the aft planet carrier assembly.

CROSS-REFERENCE TO RELATED APPLICATIONS

This application is related to concurrently filed and co-pending U.S.patent application Ser. No. ______ entitled “EPICYCLICAL GEAR SYSTEMHOUSING ASSEMBLY,” bearing Attorney Docket Number G2640-00402/RCA12398;U.S. patent application Ser. No. ______ entitled “BEARING SPRING FOREPICYCLICAL GEAR SYSTEM HOUSING ASSEMBLY,” bearing Attorney DocketNumber G2640-00404/RCA12400 and U.S. patent application Ser. No. ______entitled “STATIC CURVIC JOINT FOR EPICYCLICAL GEAR SYSTEM HOUSINGASSEMBLY,” bearing Attorney Docket Number G2640-00408/RCA12402, theentirety of each of which are herein incorporated by reference.

BACKGROUND

Epicyclical gear systems may be used in rotating machinery to transferenergy from one component, such as a rotatable shaft, to another. Byaltering certain variables such as the number, size, and teeth of thegears, an epicyclical gear system may be designed to transfer energybetween components at a desired ratio and often convert a high-speed,low-torque input to a lower-speed, higher-torque output.

Epicyclical gear systems may be suitable for a wide range ofapplications, including the transfer of energy from a turbine shaft to afan rotor in a geared turbofan engine. However, in such dynamicapplications the epicyclical gear system must be designed to allow somedegree of relative movement between parts of the system to avoidexcessive wear and, in extreme conditions, system failure.

SUMMARY

According to some aspects of the present disclosure, a planet gearhousing assembly is disclosed for an epicyclical gear system of a gasturbine engine having an engine casing. The planet gear housing assemblycomprises an aft planet carrier assembly, a forward planet carrierassembly, and a plurality of planet gears. The aft planet carrierassembly comprises an aft flange defining a central aperture and aplurality of gear shaft pockets positioned about the circumference andradially outward of the central aperture. Each pocket has a cylindricalwall, and the aft flange is coupled to the engine casing to define anaft torsional stiffness. The forward planet carrier assembly comprises aforward flange defining a central aperture and a plurality of gear shaftpockets positioned about the circumference and radially outward of thecentral aperture. Each pocket has a cylindrical wall, and the forwardflange is coupled to the aft planet carrier assembly to define a forwardtorsional stiffness. The plurality of planet gears each comprise acylindrical shaft having a forward end portion disposed in a gear shaftpocket of the forward planet carrier assembly coaxially with thecylindrical wall of the pocket, an aft end portion disposed within agear shaft pocket of the aft planet carrier assembly coaxially with thecylindrical wall of the pocket, and one or more gears carried by theshaft between the forward and aft end portions. The forward torsionalstiffness is between 60% and 80% of the aft torsional stiffness.

In some embodiments the forward torsional stiffness is between 65% and75% of the aft torsional stiffness. In some embodiments the aft planetcarrier assembly further comprises an annular mounting flange extendingfrom the aft flange, the annular mounting flange positioned forward ofand coaxial with the central aperture, the mounting flange forming aforward facing mounting surface. In some embodiments the forward planetcarrier assembly further comprises an annular mounting flange extendingfrom the forward flange, the annular mounting flange positioned aft ofand coaxial with the central aperture, the mounting flange forming anaft facing mounting surface. In some embodiments the mounting surfacesare positioned relative to each other to thereby couple the aft planetcarrier assembly and the forward planet carrier assembly.

In some embodiments the shaft and gears form a compound star gear in anepicyclical gear system. In some embodiments the forward planet carrierassembly further comprises a stiffening member positioned between theaft planet carrier assembly and the forward flange, the stiffeningmember comprising: an annular body defining a central aperture; and aplurality of radial flanges extending radially outward from the annularbody, each of the plurality of radial flanges partly defining agear-facing surface. In some embodiments the stiffening member abutsboth the forward flange and the aft planet carrier assembly.

According to further aspects of the present disclosure, a planet gearhousing assembly in an epicyclical gear assembly comprises an aft planetcarrier assembly, a forward planet carrier assembly, a plurality ofplanet gears, and a stiffening member. The aft planet carrier assemblycomprises an aft flange defining a central aperture and a plurality ofgear shaft pockets positioned about the circumference and radiallyoutward of the central aperture, each pocket having a cylindrical wall.The forward planet carrier assembly comprises a forward flange defininga central aperture and a plurality of gear shaft pockets positionedabout the circumference and radially outward of the central aperture,each pocket having a cylindrical wall. The plurality of planet gearseach comprise a cylindrical shaft having a forward end portion disposedin a gear shaft pocket of the forward planet carrier assembly coaxiallywith the cylindrical wall of the pocket, an aft end portion disposedwithin a gear shaft pocket of the aft planet carrier assembly coaxiallywith the cylindrical wall of the pocket, and one or more gears carriedby the shaft between the forward and aft end portions. The stiffeningmember is positioned between the aft planet carrier assembly and theforward planet carrier assembly. The stiffening member comprises anannular body defining a central aperture and a plurality of radialflanges extending radially outward from the annular body, each of theplurality of radial flanges partly defining a gear-facing surface.

In some embodiments the aft planet carrier assembly is coupled to theforward planet carrier assembly. In some embodiments the aft planetcarrier assembly further comprises an annular mounting flange extendingfrom the aft flange, the annular mounting flange positioned forward ofand coaxial with the central aperture, the mounting flange forming aforward facing mounting surface. In some embodiments the forward planetcarrier assembly further comprises an annular mounting flange extendingfrom the forward flange, the annular mounting flange positioned aft ofand coaxial with the central aperture, the mounting flange forming anaft facing mounting surface. In some embodiments the mounting surfacesare positioned relative to each other to thereby couple the aft planetcarrier assembly and the forward planet carrier assembly.

In some embodiments the shaft and gears form a compound star gear in anepicyclical gear system. In some embodiments the epicyclical gearassembly is a portion of a gas turbine engine having an engine casing,and wherein the aft flange is coupled to the engine casing to define anaft torsional stiffness and the forward flange is coupled to the aftplanet carrier assembly to define a forward torsional stiffness, andwherein the forward torsional stiffness is between 50% and 90% of theaft torsional stiffness. In some embodiments the forward torsionalstiffness is between 60% and 80% of the aft torsional stiffness. In someembodiments the forward torsional stiffness is between 65% and 75% ofthe aft torsional stiffness. In some embodiments the stiffening membercomprises seven gear-facing surfaces. In some embodiments the aft flangecomprises a radially outer mounting surface for mounting the planet gearhousing assembly to the engine casing.

According to still further aspects of the present disclosure, a gasturbine engine for an aircraft comprises an engine core, a fan, and agearbox. The engine core comprises a turbine, a compressor, and a coreshaft connecting the turbine to the compressor. The fan is locatedupstream of the engine core and comprises a plurality of fan blades. Thegearbox receives an input from the core shaft and outputs drive to thefan so as to drive the fan at a lower rotational speed than the coreshaft. The gearbox has a planet gear housing assembly comprising an aftplanet carrier assembly, a forward planet carrier assembly, a pluralityof planet gears, and a stiffening member. The aft planet carrierassembly comprises an aft flange defining a central aperture and aplurality of gear shaft pockets positioned about the circumference andradially outward of the central aperture, each pocket having acylindrical wall. The forward planet carrier assembly comprises aforward flange defining a central aperture and a plurality of gear shaftpockets positioned about the circumference and radially outward of thecentral aperture, each pocket having a cylindrical wall. The pluralityof planet gears each comprise a cylindrical shaft having a forward endportion disposed in a gear shaft pocket of the forward planet carrierassembly coaxially with the cylindrical wall of the pocket, an aft endportion disposed within a gear shaft pocket of the aft planet carrierassembly coaxially with the cylindrical wall of the pocket, and one ormore gears carried by the shaft between the forward and aft endportions. The stiffening member is positioned between and abutting theaft planet carrier assembly and the forward planet carrier assembly. Thestiffening member comprises an annular body defining a central apertureand a plurality of radial flanges extending radially outward from theannular body, each of the plurality of radial flanges partly defining agear-facing surface.

In some embodiments each of the plurality of planet gears comprise a sungear engaging gear and a ring gear engaging gear carried by the shaftbetween the forward and aft end portions. In some embodiments the shaftand gears form a compound star gear in an epicyclical gear system. Insome embodiments the gearbox further comprises a roller element bearingdisposed over at least a portion of one or both of the forward endportion and the aft end portion of the gear shaft. In some embodimentsthe aft planet carrier assembly defines an aft torsional stiffness andthe forward planet carrier assembly defines a forward torsionalstiffness, and wherein the forward torsional stiffness is between 65%and 75% of the aft torsional stiffness.

BRIEF DESCRIPTION OF THE DRAWINGS

The following will be apparent from elements of the figures, which areprovided for illustrative purposes.

FIG. 1 is a sectional side view of a gas turbine engine.

FIG. 2 is a close up sectional side view of an upstream portion of a gasturbine engine.

FIG. 3 is a partially cut-away view of a gearbox for a gas turbineengine.

FIG. 4 is a schematic and cross sectional view of an epicyclical gearsystem in accordance with some embodiments of the present disclosure.

FIG. 5 is a detailed schematic and cross sectional view of a planet geardisposed in an epicyclical gear system in accordance with someembodiments of the present disclosure.

FIG. 6 is an isometric view of a forward planet carrier assembly, astiffening member, and an aft carrier housing assembly of a planet gearhousing assembly in accordance with some embodiments of the presentdisclosure.

FIG. 7 is a partial cross sectional view of a stiffening member coupledbetween a forward planet carrier assembly and an aft planet carrierassembly of a planet gear housing assembly in accordance with someembodiments.

FIG. 8A is an isometric view of a stiffening member in accordance withsome embodiments.

FIG. 8B is an isometric view of a stiffening member in accordance withsome embodiments.

FIG. 9 is a flow diagram of a method in accordance with some embodimentsof the present disclosure.

The present application discloses illustrative (i.e., example)embodiments. The claimed inventions are not limited to the illustrativeembodiments. Therefore, many implementations of the claims will bedifferent than the illustrative embodiments. Various modifications canbe made to the claimed inventions without departing from the spirit andscope of the disclosure. The claims are intended to coverimplementations with such modifications.

DETAILED DESCRIPTION

For the purposes of promoting an understanding of the principles of thedisclosure, reference will now be made to a number of illustrativeembodiments in the drawings and specific language will be used todescribe the same.

As noted elsewhere herein, the present disclosure may relate to a gasturbine engine. Such a gas turbine engine may comprise an engine corecomprising a turbine, a combustor, a compressor, and a core shaftconnecting the turbine to the compressor. Such a gas turbine engine maycomprise a fan (having fan blades) located upstream of the engine core.

Arrangements of the present disclosure may be particularly, although notexclusively, beneficial for fans that are driven via a gearbox.Accordingly, the gas turbine engine may comprise a gearbox that receivesan input from the core shaft and outputs drive to the fan so as to drivethe fan at a lower rotational speed than the core shaft. The input tothe gearbox may be directly from the core shaft, or indirectly from thecore shaft, for example via a spur shaft and/or gear. The core shaft mayrigidly connect the turbine and the compressor, such that the turbineand compressor rotate at the same speed (with the fan rotating at alower speed).

The gas turbine engine as described and/or claimed herein may have anysuitable general architecture. For example, the gas turbine engine mayhave any desired number of shafts that connect turbines and compressors,for example one, two or three shafts. Purely by way of example, theturbine connected to the core shaft may be a first turbine, thecompressor connected to the core shaft may be a first compressor, andthe core shaft may be a first core shaft. The engine core may furthercomprise a second turbine, a second compressor, and a second core shaftconnecting the second turbine to the second compressor. The secondturbine, second compressor, and second core shaft may be arranged torotate at a higher rotational speed than the first core shaft.

In such an arrangement, the second compressor may be positioned axiallydownstream of the first compressor. The second compressor may bearranged to receive (for example directly receive, for example via agenerally annular duct) flow from the first compressor.

The gearbox may be arranged to be driven by the core shaft that isconfigured to rotate (for example in use) at the lowest rotational speed(for example the first core shaft in the example above). For example,the gearbox may be arranged to be driven only by the core shaft that isconfigured to rotate (for example in use) at the lowest rotational speed(for example only be the first core shaft, and not the second coreshaft, in the example above). Alternatively, the gearbox may be arrangedto be driven by any one or more shafts, for example the first and/orsecond shafts in the example above.

The gearbox may be a reduction gearbox (in that the output to the fan isa lower rotational rate than the input from the core shaft). Any type ofgearbox may be used. For example, the gearbox may be a “planetary” or“star” gearbox, as described in more detail elsewhere herein. Thegearbox may have any desired reduction ratio (defined as the rotationalspeed of the input shaft divided by the rotational speed of the outputshaft), for example greater than 2.5, for example in the range of from 3to 4.2, or 3.2 to 3.8, for example on the order of or at least 3, 3.1,3.2, 3.3, 3.4, 3.5, 3.6, 3.7, 3.8, 3.9, 4, 4.1 or 4.2. The gear ratiomay be, for example, between any two of the values in the previoussentence. Purely by way of example, the gearbox may be a “star” gearboxhaving a ratio in the range of from 3.1 or 3.2 to 3.8. In somearrangements, the gear ratio may be outside these ranges.

In any gas turbine engine as described and/or claimed herein, acombustor may be provided axially downstream of the fan andcompressor(s). For example, the combustor may be directly downstream of(for example at the exit of) the second compressor, where a secondcompressor is provided. By way of further example, the flow at the exitto the combustor may be provided to the inlet of the second turbine,where a second turbine is provided. The combustor may be providedupstream of the turbine(s).

The or each compressor (for example the first compressor and secondcompressor as described above) may comprise any number of stages, forexample multiple stages. Each stage may comprise a row of rotor bladesand a row of stator vanes, which may be variable stator vanes (in thattheir angle of incidence may be variable). The row of rotor blades andthe row of stator vanes may be axially offset from each other.

The or each turbine (for example the first turbine and second turbine asdescribed above) may comprise any number of stages, for example multiplestages. Each stage may comprise a row of rotor blades and a row ofstator vanes. The row of rotor blades and the row of stator vanes may beaxially offset from each other.

Each fan blade may be defined as having a radial span extending from aroot (or hub) at a radially inner gas-washed location, or 0% spanposition, to a tip at a 100% span position. The ratio of the radius ofthe fan blade at the hub to the radius of the fan blade at the tip maybe less than (or on the order of) any of: 0.4, 0.39, 0.38 0.37, 0.36,0.35, 0.34, 0.33, 0.32, 0.31, 0.3, 0.29, 0.28, 0.27, 0.26, or 0.25. Theratio of the radius of the fan blade at the hub to the radius of the fanblade at the tip may be in an inclusive range bounded by any two of thevalues in the previous sentence (i.e. the values may form upper or lowerbounds), for example in the range of from 0.28 to 0.32. These ratios maycommonly be referred to as the hub-to-tip ratio. The radius at the huband the radius at the tip may both be measured at the leading edge (oraxially forwardmost) part of the blade. The hub-to-tip ratio refers, ofcourse, to the gas-washed portion of the fan blade, i.e. the portionradially outside any platform.

The radius of the fan may be measured between the engine centreline andthe tip of a fan blade at its leading edge. The fan diameter (which maysimply be twice the radius of the fan) may be greater than (or on theorder of) any of: 220 cm, 230 cm, 240 cm, 250 cm (around 100 inches),260 cm, 270 cm (around 105 inches), 280 cm (around 110 inches), 290 cm(around 115 inches), 300 cm (around 120 inches), 310 cm, 320 cm (around125 inches), 330 cm (around 130 inches), 340 cm (around 135 inches), 350cm, 360 cm (around 140 inches), 370 cm (around 145 inches), 380 (around150 inches) cm, 390 cm (around 155 inches), 400 cm, 410 cm (around 160inches) or 420 cm (around 165 inches). The fan diameter may be in aninclusive range bounded by any two of the values in the previoussentence (i.e. the values may form upper or lower bounds), for examplein the range of from 240 cm to 280 cm or 330 cm to 380 cm.

The rotational speed of the fan may vary in use. Generally, therotational speed is lower for fans with a higher diameter. Purely by wayof non-limitative example, the rotational speed of the fan at cruiseconditions may be less than 2500 rpm, for example less than 2300 rpm.Purely by way of further non-limitative example, the rotational speed ofthe fan at cruise conditions for an engine having a fan diameter in therange of from 220 cm to 300 cm (for example 240 cm to 280 cm or 250 cmto 270 cm) may be in the range of from 1700 rpm to 2500 rpm, for examplein the range of from 1800 rpm to 2300 rpm, for example in the range offrom 1900 rpm to 2100 rpm. Purely by way of further non-limitativeexample, the rotational speed of the fan at cruise conditions for anengine having a fan diameter in the range of from 330 cm to 380 cm maybe in the range of from 1200 rpm to 2000 rpm, for example in the rangeof from 1300 rpm to 1800 rpm, for example in the range of from 1400 rpmto 1800 rpm.

In use of the gas turbine engine, the fan (with associated fan blades)rotates about a rotational axis. This rotation results in the tip of thefan blade moving with a velocity Utip. The work done by the fan blades13 on the flow results in an enthalpy rise dH of the flow. A fan tiploading may be defined as dH/Utip2, where dH is the enthalpy rise (forexample the 1-D average enthalpy rise) across the fan and Utip is the(translational) velocity of the fan tip, for example at the leading edgeof the tip (which may be defined as fan tip radius at leading edgemultiplied by angular speed). The fan tip loading at cruise conditionsmay be greater than (or on the order of) any of: 0.28, 0.29, 0.30, 0.31,0.32, 0.33, 0.34, 0.35, 0.36, 0.37, 0.38, 0.39 or 0.4 (all units in thisparagraph being Jkg−1K−1/(ms−1)2). The fan tip loading may be in aninclusive range bounded by any two of the values in the previoussentence (i.e. the values may form upper or lower bounds), for examplein the range of from 0.28 to 0.31, or 0.29 to 0.3.

Gas turbine engines in accordance with the present disclosure may haveany desired bypass ratio, where the bypass ratio is defined as the ratioof the mass flow rate of the flow through the bypass duct to the massflow rate of the flow through the core at cruise conditions. In somearrangements the bypass ratio may be greater than (or on the order of)any of the following: 10, 10.5, 11, 11.5, 12, 12.5, 13, 13.5, 14, 14.5,15, 15.5, 16, 16.5, 17, 17.5, 18, 18.5, 19, 19.5 or 20. The bypass ratiomay be in an inclusive range bounded by any two of the values in theprevious sentence (i.e. the values may form upper or lower bounds), forexample in the range of form 12 to 16, 13 to 15, or 13 to 14. The bypassduct may be substantially annular. The bypass duct may be radiallyoutside the core engine. The radially outer surface of the bypass ductmay be defined by a nacelle and/or a fan case.

The overall pressure ratio of a gas turbine engine as described and/orclaimed herein may be defined as the ratio of the stagnation pressureupstream of the fan to the stagnation pressure at the exit of thehighest pressure compressor (before entry into the combustor). By way ofnon-limitative example, the overall pressure ratio of a gas turbineengine as described and/or claimed herein at cruise may be greater than(or on the order of) any of the following: 35, 40, 45, 50, 55, 60, 65,70, 75. The overall pressure ratio may be in an inclusive range boundedby any two of the values in the previous sentence (i.e. the values mayform upper or lower bounds), for example in the range of from 50 to 70.

Specific thrust of an engine may be defined as the net thrust of theengine divided by the total mass flow through the engine. At cruiseconditions, the specific thrust of an engine described and/or claimedherein may be less than (or on the order of) any of the following: 110Nkg−1s, 105 Nkg−1s, 100 Nkg−1s, 95 Nkg−1s, 90 Nkg−1s, 85 Nkg−1s or 80Nkg−1s. The specific thrust may be in an inclusive range bounded by anytwo of the values in the previous sentence (i.e. the values may formupper or lower bounds), for example in the range of from 80 Nkg−1s to100 Nkg−1 s, or 85 Nkg−1 s to 95 Nkg−1 s. Such engines may beparticularly efficient in comparison with conventional gas turbineengines.

A gas turbine engine as described and/or claimed herein may have anydesired maximum thrust. Purely by way of non-limitative example, a gasturbine as described and/or claimed herein may be capable of producing amaximum thrust of at least (or on the order of) any of the following:160 kN, 170 kN, 180 kN, 190 kN, 200 kN, 250 kN, 300 kN, 350 kN, 400 kN,450 kN, 500 kN, or 550 kN. The maximum thrust may be in an inclusiverange bounded by any two of the values in the previous sentence (i.e.the values may form upper or lower bounds). Purely by way of example, agas turbine as described and/or claimed herein may be capable ofproducing a maximum thrust in the range of from 330 kN to 420 kN, forexample 350 kN to 400 kN. The thrust referred to above may be themaximum net thrust at standard atmospheric conditions at sea level plus15 degrees C. (ambient pressure 101.3 kPa, temperature 30 degrees C.),with the engine static.

In use, the temperature of the flow at the entry to the high pressureturbine may be particularly high. This temperature, which may bereferred to as TET, may be measured at the exit to the combustor, forexample immediately upstream of the first turbine vane, which itself maybe referred to as a nozzle guide vane. At cruise, the TET may be atleast (or on the order of) any of the following: 1400K, 1450K, 1500K,1550K, 1600K or 1650K. The TET at cruise may be in an inclusive rangebounded by any two of the values in the previous sentence (i.e. thevalues may form upper or lower bounds). The maximum TET in use of theengine may be, for example, at least (or on the order of) any of thefollowing: 1700K, 1750K, 1800K, 1850K, 1900K, 1950K or 2000K. Themaximum TET may be in an inclusive range bounded by any two of thevalues in the previous sentence (i.e. the values may form upper or lowerbounds), for example in the range of from 1800K to 1950K. The maximumTET may occur, for example, at a high thrust condition, for example at amaximum take-off (MTO) condition.

A fan blade and/or aerofoil portion of a fan blade described and/orclaimed herein may be manufactured from any suitable material orcombination of materials. For example at least a part of the fan bladeand/or aerofoil may be manufactured at least in part from a composite,for example a metal matrix composite and/or an organic matrix composite,such as carbon fibre. By way of further example at least a part of thefan blade and/or aerofoil may be manufactured at least in part from ametal, such as a titanium based metal or an aluminium based material(such as an aluminium-lithium alloy) or a steel based material. The fanblade may comprise at least two regions manufactured using differentmaterials. For example, the fan blade may have a protective leadingedge, which may be manufactured using a material that is better able toresist impact (for example from birds, ice or other material) than therest of the blade. Such a leading edge may, for example, be manufacturedusing titanium or a titanium-based alloy. Thus, purely by way ofexample, the fan blade may have a carbon-fibre or aluminium based body(such as an aluminium lithium alloy) with a titanium leading edge.

A fan as described and/or claimed herein may comprise a central portion,from which the fan blades may extend, for example in a radial direction.The fan blades may be attached to the central portion in any desiredmanner. For example, each fan blade may comprise a fixture which mayengage a corresponding slot in the hub (or disc). Purely by way ofexample, such a fixture may be in the form of a dovetail that may slotinto and/or engage a corresponding slot in the hub/disc in order to fixthe fan blade to the hub/disc. By way of further example, the fan bladesmaybe formed integrally with a central portion. Such an arrangement maybe referred to as a bladed disc or a bladed ring. Any suitable methodmay be used to manufacture such a bladed disc or bladed ring. Forexample, at least a part of the fan blades may be machined from a blockand/or at least part of the fan blades may be attached to the hub/discby welding, such as linear friction welding.

The gas turbine engines described and/or claimed herein may or may notbe provided with a variable area nozzle (VAN). Such a variable areanozzle may allow the exit area of the bypass duct to be varied in use.The general principles of the present disclosure may apply to engineswith or without a VAN.

The fan of a gas turbine as described and/or claimed herein may have anydesired number of fan blades, for example 14, 16, 18, 20, 22, 24 or 26fan blades.

As used herein, cruise conditions have the conventional meaning andwould be readily understood by the skilled person. Thus, for a given gasturbine engine for an aircraft, the skilled person would immediatelyrecognise cruise conditions to mean the operating point of the engine atmid-cruise of a given mission (which may be referred to in the industryas the “economic mission”) of an aircraft to which the gas turbineengine is designed to be attached. In this regard, mid-cruise is thepoint in an aircraft flight cycle at which 50% of the total fuel that isburned between top of climb and start of descent has been burned (whichmay be approximated by the midpoint—in terms of time and/ordistance—between top of climb and start of descent. Cruise conditionsthus define an operating point of the gas turbine engine that provides athrust that would ensure steady state operation (i.e. maintaining aconstant altitude and constant Mach Number) at mid-cruise of an aircraftto which it is designed to be attached, taking into account the numberof engines provided to that aircraft. For example where an engine isdesigned to be attached to an aircraft that has two engines of the sametype, at cruise conditions the engine provides half of the total thrustthat would be required for steady state operation of that aircraft atmid-cruise.

In other words, for a given gas turbine engine for an aircraft, cruiseconditions are defined as the operating point of the engine thatprovides a specified thrust (required to provide—in combination with anyother engines on the aircraft—steady state operation of the aircraft towhich it is designed to be attached at a given mid-cruise Mach Number)at the mid-cruise atmospheric conditions (defined by the InternationalStandard Atmosphere according to ISO 2533 at the mid-cruise altitude).For any given gas turbine engine for an aircraft, the mid-cruise thrust,atmospheric conditions and Mach Number are known, and thus the operatingpoint of the engine at cruise conditions is clearly defined.

Purely by way of example, the forward speed at the cruise condition maybe any point in the range of from Mach 0.7 to 0.9, for example 0.75 to0.85, for example 0.76 to 0.84, for example 0.77 to 0.83, for example0.78 to 0.82, for example 0.79 to 0.81, for example on the order of Mach0.8, on the order of Mach 0.85 or in the range of from 0.8 to 0.85. Anysingle speed within these ranges may be part of the cruise condition.For some aircraft, the cruise conditions may be outside these ranges,for example below Mach 0.7 or above Mach 0.9.

Purely by way of example, the cruise conditions may correspond tostandard atmospheric conditions (according to the International StandardAtmosphere, ISA) at an altitude that is in the range of from 10000 m to15000 m, for example in the range of from 10000 m to 12000 m, forexample in the range of from 10400 m to 11600 m (around 38000 ft), forexample in the range of from 10500 m to 11500 m, for example in therange of from 10600 m to 11400 m, for example in the range of from 10700m (around 35000 ft) to 11300 m, for example in the range of from 10800 mto 11200 m, for example in the range of from 10900 m to 11100 m, forexample on the order of 11000 m. The cruise conditions may correspond tostandard atmospheric conditions at any given altitude in these ranges.

Purely by way of example, the cruise conditions may correspond to anoperating point of the engine that provides a known required thrustlevel (for example a value in the range of from 30 kN to 35 kN) at aforward Mach number of 0.8 and standard atmospheric conditions(according to the International Standard Atmosphere) at an altitude of38000 ft (11582 m). Purely by way of further example, the cruiseconditions may correspond to an operating point of the engine thatprovides a known required thrust level (for example a value in the rangeof from 50 kN to 65 kN) at a forward Mach number of 0.85 and standardatmospheric conditions (according to the International StandardAtmosphere) at an altitude of 35000 ft (10668 m).

In use, a gas turbine engine described and/or claimed herein may operateat the cruise conditions defined elsewhere herein. Such cruiseconditions may be determined by the cruise conditions (for example themid-cruise conditions) of an aircraft to which at least one (for example2 or 4) gas turbine engine may be mounted in order to provide propulsivethrust.

According to an aspect, there is provided an aircraft comprising a gasturbine engine as described and/or claimed herein. The aircraftaccording to this aspect is the aircraft for which the gas turbineengine has been designed to be attached. Accordingly, the cruiseconditions according to this aspect correspond to the mid-cruise of theaircraft, as defined elsewhere herein.

According to an aspect, there is provided a method of operating a gasturbine engine as described and/or claimed herein. The operation may beat the cruise conditions as defined elsewhere herein (for example interms of the thrust, atmospheric conditions and Mach Number).

According to an aspect, there is provided a method of operating anaircraft comprising a gas turbine engine as described and/or claimedherein. The operation according to this aspect may include (or may be)operation at the mid-cruise of the aircraft, as defined elsewhereherein.

FIG. 1 illustrates a gas turbine engine 10 having a principal rotationalaxis 9. The engine 10 comprises an air intake 12 and a propulsive fan 23that generates two airflows: a core airflow A and a bypass airflow B.The gas turbine engine 10 comprises a core 11 that receives the coreairflow A. The engine core 11 comprises, in axial flow series, a lowpressure compressor 14, a high-pressure compressor 15, combustionequipment 16, a high-pressure turbine 17, a low pressure turbine 19 anda core exhaust nozzle 20. A nacelle 21 surrounds the gas turbine engine10 and defines a bypass duct 22 and a bypass exhaust nozzle 18. Thebypass airflow B flows through the bypass duct 22. The fan 23 isattached to and driven by the low pressure turbine 19 via a shaft 26 andan epicyclic gearbox 30.

In use, the core airflow A is accelerated and compressed by the lowpressure compressor 14 and directed into the high pressure compressor 15where further compression takes place. The compressed air exhausted fromthe high pressure compressor 15 is directed into the combustionequipment 16 where it is mixed with fuel and the mixture is combusted.The resultant hot combustion products then expand through, and therebydrive, the high pressure and low pressure turbines 17, 19 before beingexhausted through the nozzle 20 to provide some propulsive thrust. Thehigh pressure turbine 17 drives the high pressure compressor 15 by asuitable interconnecting shaft 27. The fan 23 generally provides themajority of the propulsive thrust. The epicyclic gearbox 30 is areduction gearbox.

An exemplary arrangement for a geared fan gas turbine engine 10 is shownin FIG. 2. The low pressure turbine 19 (see FIG. 1) drives the shaft 26,which is coupled to a sun wheel, or sun gear, 28 of the epicyclic geararrangement 30. Radially outwardly of the sun gear 28 and intermeshingtherewith is a plurality of planet gears 32 that are coupled together bya planet carrier 34. The planet carrier 34 constrains the planet gears32 to precess around the sun gear 28 in synchronicity whilst enablingeach planet gear 32 to rotate about its own axis. The planet carrier 34is coupled via linkages 36 to the fan 23 in order to drive its rotationabout the engine axis 9. Radially outwardly of the planet gears 32 andintermeshing therewith is an annulus or ring gear 38 that is coupled,via linkages 40, to a stationary supporting structure 24.

Note that the terms “low pressure turbine” and “low pressure compressor”as used herein may be taken to mean the lowest pressure turbine stagesand lowest pressure compressor stages (i.e. not including the fan 23)respectively and/or the turbine and compressor stages that are connectedtogether by the interconnecting shaft 26 with the lowest rotationalspeed in the engine (i.e. not including the gearbox output shaft thatdrives the fan 23). In some literature, the “low pressure turbine” and“low pressure compressor” referred to herein may alternatively be knownas the “intermediate pressure turbine” and “intermediate pressurecompressor”. Where such alternative nomenclature is used, the fan 23 maybe referred to as a first, or lowest pressure, compression stage.

The epicyclic gearbox 30 is shown by way of example in greater detail inFIG. 3. Each of the sun gear 28, planet gears 32 and ring gear 38comprise teeth about their periphery to intermesh with the other gears.However, for clarity only exemplary portions of the teeth areillustrated in FIG. 3. There are four planet gears 32 illustrated,although it will be apparent to the skilled reader that more or fewerplanet gears 32 may be provided within the scope of the claimedinvention. Practical applications of a planetary epicyclic gearbox 30generally comprise at least three planet gears 32.

The epicyclic gearbox 30 illustrated by way of example in FIGS. 2 and 3is of the planetary type, in that the planet carrier 34 is coupled to anoutput shaft via linkages 36, with the ring gear 38 fixed. However, anyother suitable type of epicyclic gearbox 30 may be used. By way offurther example, the epicyclic gearbox 30 may be a star arrangement, inwhich the planet carrier 34 is held fixed, with the ring (or annulus)gear 38 allowed to rotate. In such an arrangement the fan 23 is drivenby the ring gear 38. By way of further alternative example, the gearbox30 may be a differential gearbox in which the ring gear 38 and theplanet carrier 34 are both allowed to rotate.

It will be appreciated that the arrangement shown in FIGS. 2 and 3 is byway of example only, and various alternatives are within the scope ofthe present disclosure. Purely by way of example, any suitablearrangement may be used for locating the gearbox 30 in the engine 10and/or for connecting the gearbox 30 to the engine 10. By way of furtherexample, the connections (such as the linkages 36, 40 in the FIG. 2example) between the gearbox 30 and other parts of the engine 10 (suchas the input shaft 26, the output shaft and the fixed structure 24) mayhave any desired degree of stiffness or flexibility. By way of furtherexample, any suitable arrangement of the bearings between rotating andstationary parts of the engine (for example between the input and outputshafts from the gearbox and the fixed structures, such as the gearboxcasing) may be used, and the disclosure is not limited to the exemplaryarrangement of FIG. 2. For example, where the gearbox 30 has a stararrangement (described above), the skilled person would readilyunderstand that the arrangement of output and support linkages andbearing locations would typically be different to that shown by way ofexample in FIG. 2.

Accordingly, the present disclosure extends to a gas turbine enginehaving any arrangement of gearbox styles (for example star orplanetary), support structures, input and output shaft arrangement, andbearing locations.

Optionally, the gearbox may drive additional and/or alternativecomponents (e.g. the intermediate pressure compressor and/or a boostercompressor).

Other gas turbine engines to which the present disclosure may be appliedmay have alternative configurations. For example, such engines may havean alternative number of compressors and/or turbines and/or analternative number of interconnecting shafts. By way of further example,the gas turbine engine shown in FIG. 1 has a split flow nozzle 18, 20meaning that the flow through the bypass duct 22 has its own nozzle 18that is separate to and radially outside the core engine nozzle 20.However, this is not limiting, and any aspect of the present disclosuremay also apply to engines in which the flow through the bypass duct 22and the flow through the core 11 are mixed, or combined, before (orupstream of) a single nozzle, which may be referred to as a mixed flownozzle. One or both nozzles (whether mixed or split flow) may have afixed or variable area.

The geometry of the gas turbine engine 10, and components thereof, isdefined by a conventional axis system, comprising an axial direction(which is aligned with the rotational axis 9), a radial direction (inthe bottom-to-top direction in FIG. 1), and a circumferential direction(perpendicular to the page in the FIG. 1 view). The axial, radial andcircumferential directions are mutually perpendicular.

FIG. 4 provides a schematic view of an epicyclical gear system 100 inaccordance with some embodiments of the present disclosure. Theepicyclical gear system 100 may be a compound star gear system. A sungear 101 is coupled to and driven by a first rotatable shaft 103. Thesun gear 101 is engaged with one or more planet gears 105, such thatrotation of the sun gear 101 causes rotation of the one or more planetgears 105. The planet gears 105 may be star gears, such that the planetgears 105 rotate about an axis that is fixed relative to the axis ofrotation of the sun gear 101.

Each of the one or more planet gears 105 are engaged with a ring gear107. The ring gear 107 is coupled via a ring gear hub 108 to a secondrotatable shaft 109. Rotation of the first rotatable shaft 103 thusdrives rotation of the second rotatable shaft 109 via rotation of thesun gear 101, one or more planet gears 105, and ring gear 107. In someembodiments, the first rotatable shaft 103 may be a turbine shaft (i.e.high or low speed spool) of a turbine engine, and the second rotatableshaft 109 may be a fan shaft or fan rotor.

FIG. 5 provides a detailed and schematic view of a housing assembly 111of an epicyclical gear system 100 in accordance with some embodiments ofthe present disclosure. Each of the one or more planet gears 105 maycomprise a cylindrical gear shaft 123, a sun gear engaging gear 125, anda ring gear engaging gear 127. The sun gear engaging gear 125 and ringgear engaging gear 127 may be carried by the cylindrical gear shaft 123.Each of the one or more planet gears 105 is carried by the housingassembly 111. The planet gears 105 may be a compound star gear of theepicyclical gear system 100.

The housing assembly 111 may comprise a forward housing member 113 andan aft housing member 115. In some embodiments, the housing assembly 111further comprises an intermediate housing member 114. One or more of thehousing members 113, 114, 115 may be joined together. The forwardhousing member 113 and aft housing member 115 each define a plurality ofgear shaft pockets 116 having a cylindrical wall 120. The intermediatehousing member 114 may define a plurality of bores 118.

A cylindrical gear shaft 123 of a planet gear 105 may have a forward endportion 141 disposed within one of the plurality of gear shaft pockets116 formed by the forward housing member 113. The cylindrical gear shaft123 of the same planet gear 105 may have an aft end portion 142 disposedwithin one of the plurality of gear shaft pockets 116 formed by the afthousing member 115. The cylindrical gear shaft 123 may be disposedwithin each gear shaft pocket 116 coaxially with the cylindrical wall120 of the gear shaft pocket 116. The cylindrical gear shaft 123 mayextend through the bore 118 defined by the intermediate housing member114. The sun gear engaging gear 125 and ring gear engaging gear 127 ofplanet gear 105 may be carried by the gear shaft 123 between the forwardend portion 141 and the aft end portion 142.

The housing assembly 111 may further comprise a bearing assembly 117.The housing assembly 111 may comprise a forward bearing assembly and anaft bearing assembly. The bearing assembly 117 may comprise a bearing.For example, a forward bearing 119 may be disposed over at least aportion of the forward end portion 141 of the gear shaft 123 and an aftbearing 121 may be disposed over at least a portion of the aft endportion 142 of the gear shaft 123. Each of the bearings 119, 121 mayrotatably carry the planet gear 105. Each of the bearings 119, 121 maybe a roller element bearing.

During operation of the epicyclical gear system 100 the constituentpieces of the system 100 as described above and including additionalcarrier assemblies, housings, and bearings, may move relative to oneanother. Even small changes in the relative positioning of one componentto another can have significant impacts on performance of the system100. For example, misalignment of enmeshed gear teeth and/or bearingscan cause uneven gear and/or bearing loading, and degradation or damageto gear teeth. For example, misalignment of enmeshed gear teeth and/orbearings can cause uneven gear and/or bearing loading, degradation ordamage to gear teeth, and reduction of bearing lives and bearingstability.

Of particular concern is the fore-to-aft alignment of the gear shaft 123of each planet gear 105. Since each planet gear 105 is carried by thegear shaft 123 partly disposed within gear shaft pockets 116 of aforward housing member 113 and an aft housing member 115, relativemovement or changes in relative positioning between the forward housingmember 113 and aft housing member 115 can cause misalignment of theplanet gear 105. Similarly, relative movement or changes in relativepositioning between the intermediate housing member 114 and either orboth of the forward housing member 113 and aft housing member 115 cancause misalignment of the planet gear 105.

This misalignment may in turn lead to uneven load sharing among theplanet gears 105, gear degradation, and shortened useful life of theplanet gears 105 and/or the planet bearings 119, 121. Factors that maycontribute to planet gear misalignment include unaligned forces betweenthe forward bearing 119 and aft bearing 121, manufacturing inaccuraciesfor the positions of the gear shaft pockets 116 and bore 118, ability toreassembly the forward and aft housings in the same position that theywere machined, gear tolerance, and inflexibility and/or relativestiffness between each of the housing members 113, 114, 115.

The present disclosure is therefore directed to systems and methods forimproving and maintaining planet gear alignment in an epicyclical gearsystem. More specifically, the present disclosure is directed to aplanet gear housing assembly for an epicyclical gear system having aforward planet carrier assembly, an aft planet carrier assembly, astiffening member positioned between the forward and aft planet carrierassemblies, and a plurality of planet gears each carried by the forwardand aft planet carrier assemblies. The stiffening member may achieve adesired stiffness ratio between the forward and aft planet carrierassemblies.

As shown in FIGS. 4-8, a planet gear housing assembly 111 may comprisean aft planet carrier assembly 131, a forward planet carrier assembly161, a stiffening member 171, and a plurality of planet gears 105 eachcarried by the aft planet carrier assembly 131 and forward carrierplanet assembly. FIG. 6 is an isometric view of a forward planet carrierassembly, a stiffening member, and an aft carrier housing assembly of aplanet gear housing assembly in accordance with some embodiments of thepresent disclosure. FIG. 7 is a partial cross sectional view of astiffening member coupled between a forward planet carrier assembly andan aft planet carrier assembly of a planet gear housing assembly inaccordance with some embodiments. FIGS. 8A and 8B provide isometricviews of a stiffening member in accordance with some embodiments.

An aft planet carrier assembly 131 may comprise one or both of theintermediate housing member 114 and the aft housing member 115. The aftplanet carrier assembly 131 may comprise an aft flange 136. The aftflange 136 may be the intermediate housing member 114, the aft housingmember 115, or another flange member. The aft flange 136 may comprisemore than one flange, as shown in FIG. 6 with a first aft flange 136Aand a second aft flange 136B. The aft flange 136 may define a centralaperture 132 and a plurality of gear shaft pockets 137 positioned aboutthe circumference of and radially outward of the central aperture 132.Each of the gear shaft pockets 116 may have a cylindrical wall 138. Insome embodiments the aft flange 136 may further comprise a radiallyouter mounting surface 139 for coupling the aft planet carrier assembly131 to an engine casing.

The aft planet carrier assembly 131 may further comprise an annularmounting flange 133. The annular mounting flange 133 may be positionedforward of and coaxial with the central aperture 132. The annularmounting flange 133 may extend substantially perpendicularly from theaft flange 136. The annular mounting flange 133 may form a forwardfacing mounting surface 134 that may comprise a curvic structure.

A forward planet carrier assembly 161 may comprise a forward flange 162and an annular mounting flange 143. The forward planet carrier assembly161 may be the forward housing member 113. The forward flange 162 maydefine a central aperture 144 and a plurality of gear shaft pockets 145positioned about the circumference and radially outward of the centralaperture 144. Each gear shaft pocket 145 may have a cylindrical wall146. The annular mounting flange 143 may be positioned aft of andcoaxial with the central aperture 144. The annular mounting flange 143may form an aft facing mounting surface 147 that may comprise a curvicstructure. The annular mounting flange 143 may extend substantiallyperpendicularly from the forward flange 162.

A plurality of planet gears 105 may be carried by the forward planetcarrier assembly 161 and the aft planet carrier assembly 131. Each ofthe planet gears 105 may comprise a cylindrical shaft 123 having aforward end portion 141 disposed in a gear shaft pocket 145 of theforward flange 162. The forward end portion 141 may be disposed in thegear shaft pocket 145 coaxial with the cylindrical wall 146 defininggear shaft pocket 145. The cylindrical shaft may have an aft end portion142 disposed within a gear shaft pocket 137 of the aft flange 136, andmay be disposed in the gear shaft pocket 137 coaxial with thecylindrical wall 138.

Each planet gear 105 may further comprise one or more gears carried bythe cylindrical shaft 123 between the forward end portion 141 and theaft end portion 142. The gears may be, for example, a sun gear engagingportion 125 of the planet gear 105 and/or a ring gear engaging portion127 of the planet gear 105.

The planet gear housing assembly 111 and/or the forward planet carrierassembly 161 may further comprise a stiffening member 171. Thestiffening member 171 may comprise an annular body 172 defining acentral aperture 173. The stiffening member 171 may further comprise aplurality of radial flanges 174 extending radially outward from theannular body 172. Each of the radial flanges 174 may at least partlydefine a gear-facing surface 175. In some embodiments the stiffeningmember 171 comprises seven radial flanges 174 and defines sevengear-facing surfaces 175.

The stiffening member 171 may be positioned between the aft planetcarrier assembly 131 and the forward planet carrier assembly 161. Insome embodiments the stiffening member 171 may abut one or both of theforward flange 162 and the aft flange 136. In some embodiments thestiffening member 171 is coupled to one or both of the forward flange162 and the aft flange 136 by a plurality of bolts, pins, or otherfasteners 176.

When the epicyclical gear system 100 is fully assembled, a respectiveplanet gear 105 may be carried by the forward flange 162 and the aftflange 136, and may have a gear portion such as the ring gear engagingportion 127 positioned proximate the gear-facing surface 175 of thestiffening member 171. The stiffening member 171 may be coupled betweenthe forward flange 162 and the aft flange 136 to improve torsionalstiffness of the forward planet carrier assembly 161 with reference tothe aft planet carrier assembly 131.

The epicyclical gear system 100 may be an epicyclical gear system of agas turbine engine. The gas turbine engine may comprise an engine casing177, a portion of which is illustrated in cross section at FIG. 7. Theaft flange 136 may be coupled to the engine casing 177 to define an afttorsional stiffness. The forward flange 162 may be coupled to the aftflange 136 to define a forward torsional stiffness. The forward flange162 may be coupled to the aft flange 136, for example, with the forwardfacing mounting surface 134 of the mounting flange 133 positionedrelative to the aft facing mounting surface 147 of the mounting flange143. The forward flange 162 may be coupled to the aft flange 136 by astatic curvic joint.

In some embodiments the forward torsional stiffness may be between 60%and 80% of the aft torsional stiffness. The aft torsional stiffness maybe greater due to the mounting of the aft flange 136 to the aft casing177 at a greater radius than the forward flange 162 is mounted to theaft flange 136. More broadly, the forward torsional stiffness may bebetween 50% and 90% of the aft torsional stiffness. In other embodimentsthe forward torsional stiffness may be between 65% and 75% of the afttorsional stiffness.

FIG. 9 is a flow diagram of a method 900 of reducing relative movementbetween static components of an epicyclical gear system in accordancewith some embodiments of the present disclosure. Method 900 starts atBlock 901. The steps of method 900, presented at Blocks 901 through 919,may be performed in the order presented in FIG. 9 or in another order.One or more steps of the method 900 may not be performed.

At Block 903 a static forward planet carrier assembly 161 and a staticaft planet carrier assembly 131 may be provided. The forward planetcarrier assembly 161 may comprise a forward flange 162 defining acentral aperture 144 and a plurality of gear shaft pockets 145positioned about the circumference of the central aperture 144. Theforward planet carrier assembly 161 may further comprise an annularmounting flange 143 extending from said forward flange 162, said annularmounting flange 143 positioned aft of and coaxial with said centralaperture 144. The mounting flange 143 may form an aft facing mountingsurface 147.

The aft planet carrier assembly 131 may comprise an aft flange 136defining a central aperture 132 and a plurality of gear shaft pockets137 positioned about the circumference and radially outward of saidcentral aperture 132. The aft planet carrier assembly 131 may furthercomprise an annular mounting flange 133 extending from said aft flange136, said annular mounting flange 133 positioned forward of and coaxialwith said central aperture 132. The mounting flange 133 may form aforward facing mounting surface 134.

At Block 905 a stiffening member 171 may be positioned between said aftplanet carrier assembly 131 and said forward planet carrier assembly161. The stiffening member 171 may comprise an annular body 172 defininga central aperture 173 and a plurality of radial flanges 174 extendingradially outward from said annular body 172. Each of said plurality ofradial flanges 174 may partly define a gear-facing surface 175.

At Block 907 a bearing 119, 121 may be positioned at least partly in agear shaft pocket 145, 137 of the forward planet carrier assembly 161and/or the aft planet carrier assembly 131.

At Block 909 a planet gear 105 of a plurality of planet gears 105 may bepositioned in each of the plurality of axially aligned gear shaft pocketpairs formed between the pocket 145, 137 of the forward planet carrierassembly 161 and/or the aft planet carrier assembly 131. Each planetgear 105 may comprise a cylindrical shaft 123 having a forward endportion 141 disposed in a gear shaft pocket 145 of said forward planetcarrier assembly 161 coaxially with the cylindrical wall 146 of saidpocket 145, an aft end portion 142 disposed within a gear shaft pocket137 of said aft planet carrier assembly 131 coaxially with thecylindrical wall 138 of said pocket 137. The planet gear 105 may furthercomprise one or more gears 125, 127 carried by said shaft 123 betweensaid forward and aft end portions 141, 142.

At Block 911, a portion of a cylindrical shaft 123 of a planet gear 105may be carried with said bearing 119, 121.

At Block 913 the method may further comprise positioning a mountingsurface 134 of the aft planet carrier assembly 131 relative to amounting surface 147 of the forward planet carrier assembly 161.

At Block 915 the forward planet carrier assembly 161 may be coupled tothe aft planet carrier assembly 131.

At Block 917 the cylindrical shaft 123 of each planet gear 105 may berotated. The rotation may be driven by a sun gear 101 of a compound stargear assembly 100.

Method 900 ends at Block 919.

The presently disclosed systems and methods provide numerous advantagesover prior art systems. By providing a stiffening member positionedbetween and joining a forward and aft planet carrier assembly, thedisclosed planet gear housing assembly reduces fore-to-aft misalignmentof planet gear shafts caused by relative movement between the forwardand aft planet carrier assemblies. The aft planet carrier assembly maybe rigidly mounted to the engine casing at the outer diameter of the aftflange. Since the forward flange may be mounted to the aft flange at aninner diameter, the forward-to-aft coupling of flanges may be inherentlyless stiff and therefore allow increased torsional deflection for anequal bearing reaction load.

By coupling the aft planet carrier assembly and forward planet carrierassembly with a stiffening member between them, the forward planetcarrier assembly may be held to the aft planet carrier assembly withless relative movement between the two than in a standard coupling.While typical epicyclical gear systems may have a forward torsionalstiffness of 25-40% of the aft torsional stiffness, the disclosedstiffening member provides for an increase to a forward torsionalstiffness of 60-80% of the aft torsional stiffness. Less relativemovement is advantageous during assembly as well as operation of theepicyclical gear system.

The present disclosure may be used in combination with the disclosure ofone or more of the related applications listed above. In particular, thepresent disclosure may be used in combination with “Static Curvic Jointfor an Epicyclical Gear System Housing Assembly.” A static curvic jointassures bore-to-bore alignment during machining and assembly but maydecrease the forward torsional stiffness. The stiffening memberdescribed herein may counter this decrease in forward torsionalstiffness, such that the combination of a static curvic joint andstiffening member in an epicyclical gear system is advantageous. Thecombination may assure bore-to-bore alignment during manufacture,machining, assembly, and operation.

Although examples are illustrated and described herein, embodiments arenevertheless not limited to the details shown, since variousmodifications and structural changes may be made therein by those ofordinary skill within the scope and range of equivalents of the claims.

What is claimed is:
 1. A planet gear housing assembly in an epicyclicalgear system of a gas turbine engine having an engine casing, said planetgear housing assembly comprising: an aft planet carrier assemblycomprising an aft flange defining a central aperture and a plurality ofgear shaft pockets positioned about the circumference and radiallyoutward of said central aperture, each pocket having a cylindrical wall,wherein said aft flange is coupled to said engine casing to define anaft torsional stiffness; a forward planet carrier assembly comprising aforward flange defining a central aperture and a plurality of gear shaftpockets positioned about the circumference and radially outward of saidcentral aperture, each pocket having a cylindrical wall, wherein saidforward flange is coupled to said aft planet carrier assembly to definea forward torsional stiffness; and a plurality of planet gears eachcomprising a cylindrical shaft having a forward end portion disposed ina gear shaft pocket of said forward planet carrier assembly coaxiallywith the cylindrical wall of said pocket, an aft end portion disposedwithin a gear shaft pocket of said aft planet carrier assembly coaxiallywith the cylindrical wall of said pocket, and one or more gears carriedby said shaft between said forward and aft end portions; wherein saidforward torsional stiffness is between 60% and 80% of the aft torsionalstiffness.
 2. The planet gear housing assembly of claim 1 wherein saidforward torsional stiffness is between 65% and 75% of the aft torsionalstiffness.
 3. The planet gear housing assembly of claim 1: wherein saidaft planet carrier assembly further comprises an annular mounting flangeextending from said aft flange, said annular mounting flange positionedforward of and coaxial with said central aperture, said mounting flangeforming a forward facing mounting surface; wherein said forward planetcarrier assembly further comprises an annular mounting flange extendingfrom said forward flange, said annular mounting flange positioned aft ofand coaxial with said central aperture, said mounting flange forming anaft facing mounting surface; and wherein said mounting surfaces arepositioned relative to each other to thereby couple said aft planetcarrier assembly and said forward planet carrier assembly.
 4. The planetgear housing assembly of claim 1 wherein said shaft and gears form acompound star gear in an epicyclical gear system.
 5. The planet gearhousing assembly of claim 1 wherein said forward planet carrier assemblyfurther comprises: a stiffening member positioned between said aftplanet carrier assembly and said forward flange, said stiffening membercomprising: an annular body defining a central aperture; and a pluralityof radial flanges extending radially outward from said annular body,each of said plurality of radial flanges partly defining a gear-facingsurface.
 6. The planet gear housing assembly of claim 1 wherein saidstiffening member abuts both the forward flange and the aft planetcarrier assembly.
 7. A planet gear housing assembly in an epicyclicalgear assembly comprising: an aft planet carrier assembly comprising anaft flange defining a central aperture and a plurality of gear shaftpockets positioned about the circumference and radially outward of saidcentral aperture, each pocket having a cylindrical wall; a forwardplanet carrier assembly comprising a forward flange defining a centralaperture and a plurality of gear shaft pockets positioned about thecircumference and radially outward of said central aperture, each pockethaving a cylindrical wall; a plurality of planet gears each comprising acylindrical shaft having a forward end portion disposed in a gear shaftpocket of said forward planet carrier assembly coaxially with thecylindrical wall of said pocket, an aft end portion disposed within agear shaft pocket of said aft planet carrier assembly coaxially with thecylindrical wall of said pocket, and one or more gears carried by saidshaft between said forward and aft end portions; and a stiffening memberpositioned between said aft planet carrier assembly and said forwardplanet carrier assembly, said stiffening member comprising: an annularbody defining a central aperture; and a plurality of radial flangesextending radially outward from said annular body, each of saidplurality of radial flanges partly defining a gear-facing surface. 8.The planet gear housing assembly of claim 7 wherein said aft planetcarrier assembly is coupled to said forward planet carrier assembly. 9.The planet gear housing assembly of claim 8: wherein said aft planetcarrier assembly further comprises an annular mounting flange extendingfrom said aft flange, said annular mounting flange positioned forward ofand coaxial with said central aperture, said mounting flange forming aforward facing mounting surface; wherein said forward planet carrierassembly further comprises an annular mounting flange extending fromsaid forward flange, said annular mounting flange positioned aft of andcoaxial with said central aperture, said mounting flange forming an aftfacing mounting surface; and wherein said mounting surfaces arepositioned relative to each other to thereby couple said aft planetcarrier assembly and said forward planet carrier assembly.
 10. Theplanet gear housing assembly of claim 7 wherein said shaft and gearsform a compound star gear in an epicyclical gear system.
 11. The planetgear housing assembly of claim 7 wherein said epicyclical gear assemblyis a portion of a gas turbine engine having an engine casing, andwherein said aft flange is coupled to said engine casing to define anaft torsional stiffness and said forward flange is coupled to said aftplanet carrier assembly to define a forward torsional stiffness, andwherein said forward torsional stiffness is between 50% and 90% of theaft torsional stiffness.
 12. The planet gear housing assembly of claim11 wherein said forward torsional stiffness is between 60% and 80% ofthe aft torsional stiffness.
 13. The planet gear housing assembly ofclaim 11 wherein said forward torsional stiffness is between 65% and 75%of the aft torsional stiffness.
 14. The planet gear housing assembly ofclaim 7 wherein said stiffening member comprises seven gear-facingsurfaces.
 15. The planet gear housing assembly of claim 11 wherein saidaft flange comprises a radially outer mounting surface for mounting theplanet gear housing assembly to said engine casing.
 16. A gas turbineengine for an aircraft comprising: an engine core comprising a turbine,a compressor, and a core shaft connecting the turbine to the compressor;a fan located upstream of the engine core, the fan comprising aplurality of fan blades; and a gearbox that receives an input from thecore shaft and outputs drive to the fan so as to drive the fan at alower rotational speed than the core shaft, wherein the gearbox has aplanet gear housing assembly comprising: an aft planet carrier assemblycomprising an aft flange defining a central aperture and a plurality ofgear shaft pockets positioned about the circumference and radiallyoutward of said central aperture, each pocket having a cylindrical wall;a forward planet carrier assembly comprising a forward flange defining acentral aperture and a plurality of gear shaft pockets positioned aboutthe circumference and radially outward of said central aperture, eachpocket having a cylindrical wall; a plurality of planet gears eachcomprising a cylindrical shaft having a forward end portion disposed ina gear shaft pocket of said forward planet carrier assembly coaxiallywith the cylindrical wall of said pocket, an aft end portion disposedwithin a gear shaft pocket of said aft planet carrier assembly coaxiallywith the cylindrical wall of said pocket, and one or more gears carriedby said shaft between said forward and aft end portions; and astiffening member positioned between and abutting said aft planetcarrier assembly and said forward planet carrier assembly, saidstiffening member comprising an annular body defining a central apertureand a plurality of radial flanges extending radially outward from saidannular body, each of said plurality of radial flanges partly defining agear-facing surface.
 17. The gas turbine engine of claim 16 wherein eachof said plurality of planet gears comprise a sun gear engaging gear anda ring gear engaging gear carried by said shaft between said forward andaft end portions.
 18. The gas turbine engine of claim 16 wherein saidshaft and gears form a compound star gear in an epicyclical gear system.19. The gas turbine engine of claim 16 wherein said gearbox furthercomprises a roller element bearing disposed over at least a portion ofone or both of the forward end portion and the aft end portion of saidgear shaft.
 20. The gas turbine engine of claim 16 wherein said aftplanet carrier assembly defines an aft torsional stiffness and saidforward planet carrier assembly defines a forward torsional stiffness,and wherein said forward torsional stiffness is between 65% and 75% ofthe aft torsional stiffness.